Solar arrays contain large numbers of photovoltaic cells, also called solar cells, that convert energy from incident sunlight into electricity. A major application for the array is for spacecraft in which the solar array is used to recharge the spacecraft's DC batteries. On the array, solar cells are placed on the surface of the array, and are arranged in strings or series of electrical circuits. Those circuits are further connected in parallel to combine the outputs of the solar cells and collectively provide the appropriate levels of current at a sufficient voltage level suitable for charging the spacecraft's batteries. To ensure that the solar array receives sunlight, position control apparatus on the spacecraft track the sun's position and, by using various means, the solar arrays are kept in a position that maximizes the amount of sunlight received.
For spacecraft application, the solar arrays are light weight and are constructed to be deployable. A deployable solar array is formed of a number solar panels in a series. The panels are hinged or joined together, and can be folded up alongside one another for stowage aboard the spacecraft and transport into orbit. In orbit, upon command, the series of panels is unfurled for deployment.
Existing solar panels are flat in structure or, and generally speaking, are two dimensional. As a consequence, the panels are not very rigid. As greater amounts of electrical power are required for the spacecraft, the length and width of those panels also grows in order to accommodate greater numbers of solar cells. As the size of the array increases, the ability of the relatively thin panels to withstand the mechanical loads required to maintain adequate sun tracking decreases. Larger size panels may bend or warp, even under low gravity conditions in which the panels are essentially weightless. As the size of the array increases its inertia also increases. Although it is possible to add structural members, such as booms, to increase the array's stiffness and maintain the straightness of the array, such a modification invariably increases both the array's launch weight and its stowed volume.
Typically, the allowable stowed volume is restricted by the size of the space craft and the launch vehicle. Weight is also an important consideration in spacecraft solar array design. The higher the weight of the solar array, the more fuel is required to launch the spacecraft into orbit. The foregoing factors translate to greater launch cost, which is clearly undesirable.
The solar array's weight efficiency is therefore recognized as an important design factor. For a solar array the weight efficiency is expressed in terms of its power to weight ratio. Typical solar arrays have power to weight ratios ranging from about twenty to fifty watts per kilogram. At such low efficiencies, raising spacecraft power levels using extensions of existing solar array designs carries significant weight penalties.
The present invention has three main objectives. A principal object of the present invention therefore is to increase the power to weight ratio of solar arrays.
A further object of the invention is to enhance electrical power generation on board spacecraft, with minimal or no requirement for increasing launch weight.
An ancillary object of the invention is to increase the moment of inertia and the rigidity of solar arrays in order to increase the array's ability to withstand on orbit maneuvering loads.